1. Field of the Invention
The present invention relates generally to a gas turbine engine and more specifically to a turbine rotor blade with blade tip cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
Turbine rotor blades rotate within the engine casing and therefore form a gap with a stationary part of the turbine such as a blade outer air seal (BOAS). As the turbine heats up or cools down, the blade tip gap can change from positive to a negative value. The negative value for the blade tip gap is when tip rubbing occurs. The blade tip gap allows for hot gas to leak through and over the blade tip. The leakage flow exposes the blade tip to high temperature that can cause erosion and thus decreased turbine performance and shorter life for the blades. Blade tips are thus cooled using cooling air from the internal blade cooling circuit to limit damage to the blade tip from the high temperatures.
In the prior art, blade tip cooling is produced by drilling holes into the upper ends of the serpentine flow cooling circuit from both the pressure and suction surfaces near to the blade tip edge and the top surface of a squealer cavity. Film cooling holes are formed along the airfoil pressure side and suction side tip sections and extend from the leading edge to the trailing edge to provide cooling of the squealer tip rails. Since the blade tip region is subject to severe secondary flow field, a large number of film cooling holes and a large amount of cooling air flow are required to adequately cool the blade tip periphery. FIG. 1 shows a prior art turbine rotor blade tip with a squealer pocket and the secondary flow and cooling air pattern that forms. FIG. 2 shows the prior art blade with a row of pressure side tip periphery film cooling holes. FIG. 3 shows the film hole break-out shape for the row of film holes in FIG. 2. FIG. 4 shows the prior art blade with a row of suction side tip periphery film cooling holes.
The blade squealer tip rails are subject to heating from three exposed sides: heat load from the airfoil hot gas side surface of the tip rail; heat load from the top portion of the tip rail; and heat load from the back side of the tip rail. Cooling of the squealer pocket is performed by film cooling holes along the blade pressure and suction side periphery and conduction through the base region of the squealer pocket becomes ineffective. This is primarily due to the combination of squealer pocket geometry and the interaction of the hot gas secondary flow mixing. Thus, the effectiveness of the pressure side film cooling and tip section convective cooling holes is very limited. In addition, a TBC is normally used in the industrial engine turbine blades in order to reduce the blade metal temperature. However, applying the TBC around the blade tip rails without effective backside convection cooling may not reduce the blade tip rail metal temperature. FIG. 5 shows a prior art blade tip section cooling design with a cooling air supply channel 11, pressure side film cooling holes 12, suction side film cooling holes 13, tip rails that form a squealer pocket 14, and tip cooling holes 15 that open into the squealer pocket 14. A thermal barrier coating (TBC) 16 is applied on the walls and the tip floor. A high temperature turbine blade tip section heat load is a function of the blade tip leakage flow. A high leakage flow will induce a high heat load onto the blade tip section. Therefore, the blade tip section sealing and cooling must be addressed as a single problem.